Turbo-compressor drive for jet power plant



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632 H. 6422A W4 Y United States Patent Office 3,260,044 Patented July 12, 1966 3,260,044 TURBO-COMPRESSOR DRIVE FOR JET POWER PLANT George H. Garraway, 20 Thurman Lane, Huntington, N.Y.

Filed July 2, 1962, Ser. No. 206,735 6 Claims. (Cl. 6035.6)

This application is a continuation-in-part of copending application Serial No. 768,497, filed October 20, 1958, now Patent No. 3,164,955, dated January 12, 1965.

This invention has to do with a jet propulsion power plant system having a mechanical air compressor driven by a turbine, and provides a novel arrangement of the basic component parts of such a system. In such power plants, which develop a jet thrust by rearward discharge of an expanding stream of gas (commonly gaseous combustion products) and in which air for combustion or for the jet is brought to a high pressure by a rotary com pressor, with or without initial compression in a ram or diffuser at the air inlet, it has been common practice to employ some part or all of the high temperaturepressure gaseous products of the combustion step as the driving fluid for the turbine which drives the compressor, the gas discharged from the turbine being delivered to the jet nozzle at a super atmospheric pressure. In proposals for using nuclear energy in such jet propulsion power plants, the step of air-fuel combustion is replaced by an air heating step and the heated air, precompressed, is to form the jet. For the present purpose, the two are equivalent and are referred to generally as the step of providing a hot gaseous jet fluid.

The practice of using some or all of the high temperature and pressure jet fluid to drive the turbine has imposed a limitation on the heating or combustion process because the temperature of the heated air or of the combustion products passing through the turbine must be limited to a value tolerable by the metal of the turbine. This limitation requires that one or more of all of the factors affecting the temperature of the gaseous jet fluid be limited, such as the degree of compression of the air (in a combustion system) or the fuel/air ratio or the kind of fuel.

There has been a diificulty in finding the best compromise of these factors while working within the imposed limitation upon the attained temperature. Such practices sometimes lower the efliciency of the combustion and always decrease the propulsive force of the jet because of the limitation upon the temperature and pressure of the gaseous fluid delivered to the jet. Further, the use of the gaseous jet fluid as the turbine drive fluid has involved the diversion of a large part of the thus limited energy potential of that fluid. The elfort to minimize the effect of these limitations has been directed to improvements in turbine and compressor design to increase efliciency, and to improvement of metals to permit higher gas temperatures through the turbine; but a serious temperature limitation has remained.

In my copending application Serial No. 768,497, filed October 20, 1958, now Patent No. 3,164,955, dated January 12, 1965, a system is described which obviates the aforementional limitations by driving the turbine of the turbo-compressor unit by a separate fluid which is in a gaseous state at least when passing through the turbine and which is circulated in a closed loop including a heat exchange stage at which the circulating fluid is heated by indirect heat exchange with a hot jet fluid at one or more points in the path from air inlet to jet discharge. In such a system, the selections of the fuel, the degree of compression and such factors and the fuel/air ratio for combustion, and also of the temperatures and pressures attained in the gaseous jet fluid, are independent of any limitations imposed by the turbine, and therefore are such as to give a greater energy drop through the jet nozzle and a greater propulsive force from the jet, often with improved heating or combustion efliciency.

The present invention accomplishes its object of improving the operation of such a jet power plant system by a novel arrangement of air flow through compressor, turbine, and heating chamber, which permits the attainment of higher temperatures in a high pressure and temperature air heating or combustion step (of a large portion of the air), and consequently a greater jet thrust. The improved performance is evidenced under take-off conditions, but to an even greater extent at higher speeds and altitude with either a decrease in initial fuel load or an increase in distance of travel or increased speed for the same fuel load or both. All in comparison with a system in which the gaseous products generated in the high temperature-pressure heating step are used as the turbine drive fluid.

The present invention accomplishes the major objectives attained by the system described in Serial No. 768,- 497, in a more simplified manner and without the necessity of providing a closed loop turbinedriving secondary fluid and the concomitant storage problems and complexities related thereto.

The system described in Serial No. 768,497, through the use of a completely separate turbine driving fluid, allows for the utilization of the entire air mass discharged from the compressor in a high pressure heating chamber. The system of the present invention utilizes a substantial portion of the air discharged from the compressor in a high temperature-pressure heating chamber (main) with however a portion of the air from the compressor diverted for expansion through the turbine to a lower pressure before being utilized in a relatively low temperature-pressure heating chamber (secondary). While the combined thrust potential in the system of this invention is somewhat less than the thrust potential resulting from the system in Serial No. 768,497, it is a substantial improvement as compared to that produced by existing conventional designs. There are advantages in this modified design which tend to compensate for a slightly lowered thrust. It is less complex in that it eliminates the necessity of a separate turbine driving fluid circuit, yet it retains the advantage of being available to mass cool the high pressure heating chamber and jet nozzle and/or the air discharging from the ram diffuser before entry into the compressor. Furthermore, turbine design for expansion of a portion of the compressed air is similar to that for existing direct air cycle engines, and could be relatively simply adapted to use through modification of existing designs or use of presently available engine components by those skilled in the art having the benefit of this disclosure.

In the system of this invention the turbine of the turbocornpressor unit is driven by diverting a portion of the air mass flowing from the compressor and expanding same through the turbine. A further feature of the invention is that after expansion through the turbine, that portion of the total air mass being new at a relatively low pressure and temperature, may be passed through a heat exchange stage and heated by indirect heat exchange with a hot jet fluid at one or more points in the path from the air inlet to jet discharge. The turbine driving fluid may be used as a mass coolant at the high temperature and pressure heating or combustion step, at the jet nozzle or to cool the hot air passing from the ram diffuser to the compressor at high speed or any combination of these.

The portion of air discharging from the compressor which is not diverted to expand through the turbine, is

passed directly into the main heating chamber wherein it may be heated to an optimum temperature and pressure and then expanded directly through the jet nozzle to thereby provide thrust.

In this system, as in my copending application referred to above, the selections of the fuel, the degree of compression and such factors as the fuel/air ratio for combustion of that portion of the air which is not expanded through the turbine, and also of the temperatures and pressures attained in the gaseous jet fluid, are independent of any limitations imposed by the turbine, and therefore are such as to give a greater pressure and temperatur drop through the jet nozzle and a greater propulsive force from the jet, often with improved heating or combustion efliciency. Also, it becomes unnecessary to use such a heavy multi-stage rotary compressor, and a simpler and lighter compressor may be used which eflects a reduction of weight and cost as compared with prior art systems.

Other advantages, and particular preferred features, are described below in the description of an illustrative form shown schematically in FIGURE 1 of the accompanying drawings. FIGURE 2 shows a similar system schematically, including a further feature of novelty. FIGURE 3 shows a schematic diagram of a coaxial arrangement of the main combustor-jet nozzle and the secondary combustor-jet nozzle. FIGURE 4 is a schematic diagram illustrating a special coaxial arrangement of the respective main and secondary eombustors and jet nozzles.

The illustrative system of the FIGURE 1 diagram is one in which the gaseous jet fluid consists of products of fuel-air combustion, but it will be understood that the same principle is applicable when that fluid is air heated by nuclear energy. In this system air is taken in for initial compression in a conventional ram or diffuser step 10, which may comprise one or more diffuser units and which delivers to the mechanical rotary compressor 11 driven in any suitable way (indicated by the dot-dash line 12) by the rotary gas turbine 13.

The cornpressed air product of the compressor 11 is divided into two streams, 14 and 15, respectively. Stream '15, which represents a portion of the compressed air from compressor 11, is delivered to a main combustor 16 which in turn delivers to a jet nozzle 17. When nuclear energy is used, either alone or with airfuel combustion systems, the combustor 16 is either replaced or supplemented by a heater in which the pre-compressed air from compressor 11 is heated by the source of nuclear energy. The construction of these several parts is not a part of the present novelty and may take any of various forms. When air-fuel combustion is used, fuel in fluid state is supplied to the combustor 16 by a line 18 which may include one or more heat exchangers (not shown) for preheating and preferably vaporizing the fuel before it is delivered to the combustor.

Stream 14 is the portion of compressed air from compressor 11 delivered to the turbine 13 through heater 28, for expansion therethrough, thereby operating the turbine 13. The air is delivered from the turbine 13 at a relatively low temperature and pressure to a secondary combustor 19, wherein it is admixed with fuel delivered to the combustor 19 from line 20. The gases exiting from combustor 19 are passed through the secondary jet nozzle 21.

A heating zone 28, in line 14 for elevating the temperature of the air discharging from the compressor 11 prior to passing a portion of said air to the turbine 13, is disposed in the compressor discharge line 14. The heating zone may be a combustion zone wherein the temperature of the exiting gases is maintained at or below that tolerable at the turbine inlet. While the heating zone has been shown in FIGURE 1 to be disposed in line 14 and the air delivered to the combustor 16 unheated, it is to be understood that said heating zone may likewise be disposed in the compressor discharge line prior to division of the compressor discharge air into lines 14 and 15.

The turbine 13 may be coaxial with the compressor 11 or may be on a separate axis, with any suitable mechanical driving connection between the two in either case, various designs for this purpose being well known. The turbine 13 may also serve to drive auxiliaries of the jet power plant.

One of the advantages of the system of this invention is that it permits greater freedom as to the spacing and relative location of the main combustor, the turbine and the jet nozzle, whereas in systems using gaseous jet fluid to drive the turbine it is almost imperative that the main combustor be close to the turbine 13. This greater freedom in physical design is of special value where a multiple jet nozzle is used.

The ram diffuser 10 for effecting an initial compression of the air taken in at the air inlet is not essential to the basic system but is highly desirable. Its potential advantages are more fully realized in the present system because of the possibility of conducting the combustion or heating in a way to attain higher pressure in the hot gaseous jet fluid.

In this system, requiring no resort to excess combustion air to hold down the temperature of the combustion products in a combustion system, the ratio of fuel to air can be substantially stoichiometric, and the main combustor may be so designed on known principles as to withstand combustion products of higher temperature and pressure than in the past, since the combustion products in passing to and through the jet nozzle come in contact only with static surfaces which can be mass cooled. A greater energy drop through the jet nozzle is therefore possible and a consequent greater propulsive effort or thrust.

A further contribution to the gain in thrust is available because in this system, with its higher temperature and pressure of the gaseous jet fluid, a much smaller fraction of the energy content of that fluid is diverted to turbine operation than is necessary when the jet fluid itself is used as the turbine drive fluid. Operation of the jetnozzle with jet fluid of higher temperatures than in the past is facilitated especially when the drive fluid (air) for the turbine is passed in indirect heat exchange with the nozzle and main combustor or heater 16, since this provides a good means of cooling the walls of these high temperature units. This feature is more fully explained hereafter.

If the system of the present invention is used in a jet power plant that develops the same jet thrust as a comparable plant using gaseous jet fluid to drive the turbine, the advantage of the present system is realized as a large reduction of the rate of air flow through the compressor, combustor (heater) and jet nozzle. Operation of aircraft at higher altitudes and speeds becomes possible with this system, both because the increase in jet thrust permits easier attainment of high altitudes and because, with less air available at the high altitude, the system permits attainment of a suflicient thrust to maintain flight. All of these advantages are expressed as an increase of thrust per unit weight of air flow per second due to the higher attainable pressure and temperature of the gaseous jet fluid.

The increased thrust attainable with this system makes it feasible to employ an air-breathing jet power plant as the first stage of a multi-stage missile or space vehicle. The engine Weight and fuel load required for a jet power plant so used, and employing this system, is less than that for a liquid or solid fuel rocket, so that the advantage can be realized as either a faster acceleration or a longer first-stage flight, or both, or an increase in ultimate payload, or simply as a reduction in weight for the same payload.

The advantage of greater jet thrust per unit of air flow per second appears also in the smaller percentage reduction of thrust with increase in air inlet velocity. As is known, jet thrust is a function of the difference between outlet (jet) velocity and air inlet velocity, so that thrust is reduced as the vehicle attains a higher velocity causing a higher air inlet velocity. The subtractive effect of inlet velocity, being a constant, at any stated speed is smaller in terms of the percentage reduction of thrust when the jet velocity is greater, as it can be when the pressure and/or temperature of the gaseous jet fluid delivered to the jet nozzle can be increased, as with the system of this invention.

FIGURE 2 is a similar schematic flow diagram of a system having all of the elements shown in FIGURE 1 in the same relationship as in the system of FIGURE 1, but with certain additional elements now described, the flow of air to the combustor and turbine being as shown by the arrows. Here again, the illustrative form is one using air-fuel combustion to provide the jet fluid, but the same principle is applicable if the combustor is replaced or supplemented by a heater for pre-compressed air with nuclear fuel as the source of heat.

The additions in the system of FIGURE 2 are (1) a heat exchanger 22 effecting indirect heat exchange between (a) the inflowing air discharging at elevated pressure and temperature from the ram diffuser 10, and (b) the turbine-driven air exiting from the turbine 13 through line 24; and (2) a heat exchanger 23 effecting heat exchange between (a) the outer walls of the high temperature pressure combustor 16 and jet nozzle 17, and (b) the turbine-driven air delivered from turbine 13, through line 25.

In FIGURE 2, the air entering the ram diffuser travels a path as in FIGURE 1, including dividing the air leaving the compressor 11 and delivery of one portion to the high temperature and pressure (main) combustor 16 and the other portion to the turbine 13. As in FIGURE 1, the heater 28 is employed to heat the air to be expanded through the turbine 13 during the periods of take-off, climb and acceleration. Under supersonic or hypersonic flight conditions, the air passing through the turbine 13 from the compressor 11 does not require additional heating after leaving the compressor and prior to entry into the turbine 13, and therefore the heater 28 need not be operated during this period. In the feature illustrated in FIGURE 2, the air delivered from the turbine 13 is passed to the heat exchanger 22 and then to the heat exchanger 23 before delivery to the low temperature and pressure (secondary) heat combustor 19 through lines 24, and/or 25 and 26. It should be understood that the air may be delivered from the turbine 13 through line 25, heat exchanger 23 and line 26, without passage through line 24 and heat exchanger 22, if desired.

At very high speeds of a jet propelled air-breathing aircraft, the ram diffuser 10 effects a great increase in pressure of the air passing through it, by the well known ram action. This involves a great increase in air temperature as well, a matter which imposes a limitation upon the speed at which the aircraft can operate Well because of the adverse effect upon the rotary compressor 11. Because of the increase in jet propulsive force attainable with my system of turbine drive, with its removal of the limit upon the tolerable temperature of the compressor 11 output, air speeds are attainable which would involve an excessively high temperature in the output of a ram diffuser. By placing the heat exchanger 22 at the ram diffuser outlet, and ahead of the rotary mechanical compressor 11, I provide a heat sink which lowers the temperature and reduces the specific volume of the pre-compressed inlet air passing to the rotary compressor 11, and thereby the latter is relieved of what would otherwise be an undesirable or even impracticable temperature condition. This reduction of specific volume at the inlet of compressor 1! permits that compressor to handle a greater weight of air flow per second. The greater weight of air flow may be provided by the same ram diffuser 10, in a case where otherwise the rotary compressor is the limiting factor, and permits the use of additional ram diffuser capacity as by the provision of multiple diffusers. Further, the reduction in temperature of the air at the inlet of the compressor results in a greater pressure rise of the air in the rotary compressor.

The heat exchanger 22 can be by-passed when desired, by closing valve 27 and opening valve 29. This is for operation at lower aircraft speeds where the heat effect of the ram action is insignificant. In that event, the turbine-driven air delivered from turbine 13 is passed directly to the heat exchanger 23 instead of first being heated in exchanger 22.

Considerations of the engineering design of a particular jet power plant may result in passing the turbine drive fluid first through heat exchanger 23 and then through heat exchanger 22. Moreover, design may dictate passing controlled portions of the turbine discharge air through both lines 24 and 25, simultaneously and then admixing same before or at combustion; I include these variants within my invention.

The invention has been illustrated showing a single stage compressor, however, if desirable, a multiple stage compressor may be used. Likewise, a multiple combustor for either the main and/ or secondary combustor products may be employed.

FIGURE 3 illustrates a schematic of a novel coaxial arrangement of a main and secondary heating chamber and respective jet nozzles. A further feature, shown in FIGURE 4, is a schematic showing an alternate positioning of the main combustor and concomitant jet nozzle in a coaxial arrangement with the secondary combustor such that there is an incremential expansion of the high pressure combustion gases to the pressure level of the turbine exhaust gases.

In FIGURE 3, a main combustor and jet nozzle 30 and 30a respectively, are disposed within a duct 31 carrying the turbine discharge air in a manner such that the turbine discharge gases may pass over the outer shell of the main combustor and thereby mass cool same. The duct 31 transitions into a secondary combustion chamber at 32 wherein fuel 33 is added (reheat or after burning) prior to expansion through the secondary annular jet nozzle 34. Note that the thrust nozzle 30a receiving the gases from the main combustor 30 is disposed coaxially within the thrust nozzle receiving gases from the secondary combustor.

FIGURE 4 shows a modification of FIGURE 3 representing a further feature of this invention. In FIGURE 4, the high pressure combustor 30 and jet nozzle 300! are disposed at a point rearward of the converging portion of the thrust nozzle 34. This arrangement provides an incremental expansion of the relatively high pressure gases from the combustor 30 and jet nozzle 30a to the pressure level of the turbine discharge air with a concomitant production of increased thrust due to reaction against the inner walls of the main combustion chamber; the high velocity gases which result are then available to mix with the low pressure turbine exhaust air moving at low velocity (mixing may be aided by utilizing known elements for efficient mixing, such as diffuser vanes, not shown, just prior to or at the point of fuel addition to this low pressure turbine exhaust air). Through an interchange of kinetic energy which is otherwise wasted, a high approach velocity is produced of the resultant mixture which subsequently through expansion from the secondary jet nozzle is further increased in velocity. The result is a greater thrust than if the high pressure and the low pressure expansion nozzles operate independently in expanding back to atmosphere pressure regardless of coaxial or separately located.

I claim:

1. A jet propulsion power system comprising an air inlet, a mechanical compressor for inlet air, a gas turbine for driving the compressor, a hot gaseous jet fluid gencrating chamber receiving a portion of the compressed air from the compressor, a jet nozzle receiving hot gaseous jet fluid from the generating chamber and forming a propulsive jet, means for supply a driving fluid to said turbine in a gaseous state which comprises a flow conduit between the compressor discharge and the turbine inlet for delivering the remaining portion of air from the compressor to the turbine inlet, and means therein for elevating the temperature of the air passing to the turbine inlet, said chamber and said nozzle being comprised of a unitary structure, said structure having an unbroken exterior surface continuous over said chamber and nozzle portion thereof, and a flow conduit from the turbine discharge arranged to conduct turbine discharge fluid about said unitary structure to effect indirect heat exchange between turbine discharge fluid and relatively hot jet fluid Whereby said structure is thereby capable of being sufficiently cooled by mass flow of turbine discharge fluid to generate said jet fluid in said chamber by stoichiometric combustion and to conduct through said nozzle said jet fluid generated thereby.

2. A system in accordance with claim 1 and further ineluding a ram diffuser for receiving and compressing inlet air ahead of said mechanical compressor and means including said turbine discharge flow conduit for effecting indirect heat exchange between said turbine driving fluid and relatively hot air discharging from said ram diffuser, said turbine discharge conduit being further arranged to conduct said turbine driving fluid in heat exchange relation with said diffuser prior to conducting said fluid about said unitary structure.

3. A system in accordance with claim 2 and further including a secondary jet fluid generating chamber for receiving turbine discharge fluid and a secondary jet nozzle for receiving the jet fluid from the secondary generating chamber, said turbine discharge flow conduit being arranged to conduct said turbine discharge fluid to said secondary chamber after said fluid passes said unitary structure.

4. In a jet propulsion power system comprising a turbine, a turbine driven compressor for compressing inlet air and a jet fluid generating chamber receiving a first portion of the air thus compressed and delivering said fluid to a jet-forming nozzle, said chamber and said nozzle being comprised of a unitary structure having an unbroken exterior surface continuous over said chamber and nozzle portion thereof, the improvement which comprises combusting substantially stoichiometrically said first portion of the air in said chamber, delivering a motive fluid including the remaining portion of the compressed air to and through the turbine, passing the motive fluid discharged from the turbine over the exterior surface of said unitary structure thereby effecting an indirect heat exchange between the turbine discharge fluid and relatively hot jet fluid.

5. The method of claim 4 wherein the inlet air is passed through a ram diffuser ahead of the compressor and there is effected an indirect heat exchange between the turbine discharge fluid and relatively hot air discharging from said ram diffuser.

6. The method of claim 5 wherein the turbine discharge fluid after indirect heat exchange with the air discharging from the ram diffuser is passed to a secondary jet fluid generating chamber and subsequently to a secondary jet nozzle.

References Cited by the Examiner UNITED STATES PATENTS 2,577,919 12/1951 Roy 35.6 2,592,938 4/1952 McNaught 6035.6 2,603,946 7/1952 Lagelbauer 6035.6 2,712,727 7/1955 Morley et a1. 6039.51 X 2,930,190 3/1960 Rogers 603918 3,054,254 9/1962 Hopper 6035.6 3,063,241 11/1962 Langfelder 6035.55 3,102,385 9/1963 Lyons 6035.6

MARK M. NEWMAN, Primary Examiner. ABRAM BLUM, SAMUEL LEVINE, Examiners. S. N. GARBER, D. HART, Assistant Examiners. 

1. A JET PROPULSION POWER SYSTEM COMPRISING AN AIR INLET, A MECHANICAL COMPRESSOR FOR INLET AIR, A GAS TURBINE FOR DRIVING THE COMPRESSOR, A HOT GASEOUS JET FLUID GENERATING CHAMBER RECEIVING A PORTION OF THE COMPRESSED AIR FROM THE COMPRESSOR, A JET NOZZLE RECEIVING HOT GASEOUS JET FLUID FROM THE GENERATING CHAMBER AND FORMING A PROPULSIVE JET, MEANS FOR SUPPLY A DRIVING FLUID TO SAID TURBINE IN A GASEOUS STATE WHICH COMPRISES A FLOW CONDUIT BETWEEN THE COMPRESSOR DISCHARGE AND THE TURBINE INLET FOR DELIVERING THE REMAINING PORTION OF AIR FROM THE COMPRESSOR TO THE TURBINE INLET, AND MEANS THEREIN FOR ELEVATING THE TEMPERATURE OF THE AIR PASSING TO THE TURBINE INLET, SAID CHAMBER AND SAID NOZZLE BEING COMPRESSED TO A UNITARY STRUCTURE, SAID STRUCTURE HAVING AN UNBROKEN EXTERIOR SURFACE CONTINUOUS OVER SAID CHAMBER AND NOZZLE PORTION THEREOF, AND A FLOW CONDUIT FROM THE TURBINE DISCHARGE ARRANGED TO CONDUCT TURBINE DICHARGE FLUID ABOUT SAID UNITARY STRUCTURE TO EFFECT INDIRECT HEAT EXCHANGE BETWEEN TURBINE DISCHARGE FLUID AND RELATIVELY HOT JET FLUID WHEREBY SAID STRUCTURE IS THEREBY CAPABLE OF BEING SUFFICIENTLY COOLED BY MASS FLOW OF TURBINE DISCHARGE FLUID TO GENERATE SAID JET FLUID IN SAID CHAMBER BY STOICHIOMETRIC COMBUSTION AND TO CONDUCT THROUGH SAID NOZZLE SAID JET FLUID GENERATED THEREBY. 